Protective coating systems for gas turbine engine applications and methods for fabricating the same

ABSTRACT

A protective coating system includes a turbine engine component substrate formed of a ceramic matrix composite material, an environmental barrier coating layer including a rare earth disilicate material formed directly on the substrate, and a thermal barrier coating layer including a porous rare earth monosilicate material having a metal silicate material infiltrated within at least a portion of the pores formed directly on the environmental barrier coating layer.

TECHNICAL FIELD

The present disclosure generally relates to protective coatings for gasturbine engine applications and methods for fabricating such coatings.

BACKGROUND

Turbine engines are used as the primary power source for various kindsof aircraft and other vehicles. The engines may also serve as auxiliarypower sources that drive air compressors, hydraulic pumps, andindustrial electrical power generators. Most turbine engines generallyfollow the same basic power generation procedure. Compressed air ismixed with fuel and burned, and the expanding hot combustion gases aredirected against stationary and rotary turbine vanes in the engine. Thevanes turn the high velocity gas flow partially sideways to impinge ontoturbine blades mounted on a rotatable turbine disk. The force of theimpinging gas causes the turbine disk to spin at high speed. Jetpropulsion engines use the power created by the rotating turbine disk todraw more air into the engine, and the high velocity combustion gas ispassed out of the gas turbine aft end to create forward thrust. Otherengines use this power to turn one or more propellers, electricalgenerators, or other devices.

High temperature turbine components such as turbine blades, shrouds, andnozzles made from silicon nitride or silicon carbide have the potentialto appreciably increase the operating temperatures of turbine engines.The high temperature and high pressure environment of the turbine engineas well as the high gas velocity can cause erosion of silicon basedceramics. The mechanism of some of the erosion loss is due to theformation of SiO₂ and SiO gas. Typically, combustion gas environments,including turbine engines, contain about 10% water vapor. Oxygencontaining water in the turbine reacts with silicon nitride or siliconcarbide to form silica scale on silicon based ceramic surfaces. Watervapor can also react with the silica scale to form silicon hydroxide,which is volatile. Evaporation of silicon hydroxide from ceramicsurfaces and erosion of ceramic caused by high speed combustion gasespassing over ceramic surfaces leads to the loss of ceramic material fromceramic combustor and turbine components.

U.S. Pat. No. 6,159,553 and US 2002/0136835 A1 show protective ceramiccoatings. Tantalum oxide alloyed with lanthanum oxide provides anenvironmental barrier coating (EBC). However, tantalum oxide permitsdiffusion of oxygen, resulting in the formation of a SiO2 layer belowthe tantalum oxide layer. Published U.S. patent application 2002/0098391by Tanaka et al discloses the use of rare earth silicates to form aprotective coating to a silicon based substrate ceramic material. Butthe process disclosed by Tanaka limits the coating composition becauseit allows interaction of the coating with the substrate.

Accordingly, there is a need for an improved coating and method to applythe coating for a high temperature (>2600° F. (>1425° C.)) barrierbetween an environmental coating and a substrate of silicon nitride orsilicon carbide. There is also a need for a diffusion coating that willprevent migration of cations out of a silicon-based substrate. There isas well a need to coat complex parts with a uniform dense oxidationresistant coating at a minimal cost. Furthermore, other desirablefeatures and characteristics of the present invention will becomeapparent from the subsequent detailed description of the invention andthe appended claims, taken in conjunction with the accompanying drawingsand this background of the invention.

BRIEF SUMMARY

The present disclosure generally relates to protective coatings for gasturbine engine applications and methods for fabricating such coatings.In one embodiment, a protective coating system includes a turbine enginecomponent substrate formed of a ceramic matrix composite material, anenvironmental barrier coating layer including a rare earth disilicatematerial formed directly on the substrate, and a thermal barrier coatinglayer including a porous rare earth monosilicate material having a metalsilicate material infiltrated within at least a portion of the poresformed directly on the environmental barrier coating layer.

In another embodiment, a method of applying a protective coating to asubstrate includes the steps of: providing a turbine engine componentsubstrate formed of a ceramic matrix composite material, forming anenvironmental barrier coating layer including a rare earth disilicatematerial directly on the substrate, treating an outer surface of theenvironmental barrier coating layer to form a thermal barrier coatinglayer including a porous rare earth monociliate material directly on theenvironmental barrier coating layer, and infiltrating at least a portionof the pores with a metal solution or suspension.

This summary is provided to introduce a selection of concepts in asimplified form that are further described below in the detaileddescription. This summary is not intended to identify key features oressential features of the claimed subject matter, nor is it intended tobe used as an aid in determining the scope of the claimed subjectmatter.

BRIEF DESCRIPTION OF THE DRAWING

The present invention will hereinafter be described in conjunction withthe following drawing figures, wherein like numerals denote likeelements, and wherein:

FIG. 1 is a partial cross-sectional view of a gas turbine engine inaccordance with an exemplary embodiment;

FIG. 2 is a partial, sectional elevation view illustrating a portion ofa turbine section of the gas turbine engine of FIG. 1 in accordance withan exemplary embodiment;

FIG. 3 illustrates, in cross section, the surface of an exemplaryturbine component substrate made of a SiC-SiC material, in one exemplaryembodiment;

FIG. 4 illustrates, in cross section, the surface of the exemplaryturbine component substrate of FIG. 3 , having been coated with a rareearth disilicate EBC layer, in one exemplary embodiment;

FIG. 5 illustrates, in cross section, the surface of the exemplaryturbine component substrate of FIG. 3 , having been coated with a rareearth disilicate EBC layer that has been partially converted at itsouter surface to a rare earth monosilicate, in one exemplary embodiment;

FIG. 6 is a cross-sectional micrograph of a disilicate EBC layer havingbeen thermally treated to form porous monosilicate layer, in oneexemplary embodiment;

FIG. 7 is a cross-sectional micrograph of a disilicate EBC layer havingbeen chemically-thermally treated to form porous monosilicate layer, inone exemplary embodiment;

FIG. 8 is a process flow diagram for chemically-thermally converting adisilicate EBC layer to a porous monosilicate layer, in one exemplaryembodiment;

FIG. 9 is a cross-sectional micrograph of a disilicate EBC layer havingbeen thermally treated to form porous monosilicate layer and havinginfiltrated therein one or more metal salts, nitrates, carbonates oroxides, in one exemplary embodiment; and

FIG. 10 is a cross-sectional micrograph of a disilicate EBC layer havingbeen chemically-thermally treated to form porous monosilicate layer andhaving infiltrated therein one or more metal salts, nitrates, carbonatesor oxides, in one exemplary embodiment;

DETAILED DESCRIPTION

The following detailed description of the invention is merely exemplaryin nature and is not intended to limit the invention or the applicationand uses of the invention. Furthermore, there is no intention to bebound by any theory presented in the preceding background of theinvention or the following detailed description of the invention.

Silicon carbide-silicon carbide matrix (“SiC—SiC”) and silicon nitride(“Si₃N₄” or simply “SiN”) materials are currently limited in operationaluse temperature by oxidation which begins around 2400° F., or even lowerin some instances. While there are many coating methods that have beenput forth, all claiming to resolve the issues of other methods, theyeach have issues of their own. In other words, gaining a benefit in aproperty from one process or material often leads to a shortfall inanother property. The present disclosure provides an approach tocreating an oxidation/thermal barrier coating for SiC-SiC or SiNsubstrate materials to allow the use temperature to be increased toabout 2600° F. to about 2800° F. The approach employs a rare earthdisilicate EBC coating that is disposed onto the substrate, and thensubjected to thermal or carbon monoxide-based processing to decompose anouter surface of the rare earth disilicate coating to form a rare earthmonosilicate porous outer layer. This porous outer layer may then beinfiltrated with a metal salt solution or a metal oxide nanoparticlesuspension. Further thermal processing allows the metal oxide or salt(which converts to oxide) to chemically integrate with the porousmonosilicate layer, effectively forming a thermal barrier coating (TBC)layer over the rare earth disilicate EBC coating, which may increasemelting point temperature capability beyond the capability of themonosilicate.

Turbine Engine/Turbine Section

As initially noted, embodiments of the present disclosure findparticular application in the “hot” or turbine sections of gas turbineengines. Turning now to the Figures, FIG. 1 is a cross-sectional view ofa gas turbine engine 100 according to an exemplary embodiment. AlthoughFIG. 1 depicts a turbofan engine, in general, exemplary embodimentsdiscussed herein may be applicable to any type of engine, includingturboshaft engines. The gas turbine engine 100 may form part of, forexample, an auxiliary power unit for an aircraft or a propulsion systemfor an aircraft. The gas turbine engine 100 has an overall constructionand operation that is generally understood by persons skilled in theart. The gas turbine engine 100 may be disposed in an engine case 101and may include a fan section 120, a compressor section 130, acombustion section 140, a turbine section 150, and an exhaust section160. The fan section 120 may include a fan, which draws in andaccelerates air. A fraction of the accelerated air from the fan section120 is directed through a bypass section 170 to provide a forwardthrust. The remaining fraction of air exhausted from the fan is directedinto the compressor section 130.

The compressor section 130 may include a series of compressors thatraise the pressure of the air directed into it from the fan section 120.The compressors may direct the compressed air into the combustionsection 140. In the combustion section 140, the high pressure air ismixed with fuel and combusted. The combusted air is then directed intothe turbine section 150. The turbine section 150 may include a series ofrotor and stator assemblies disposed in axial flow series. The combustedair from the combustion section 140 expands through the rotor and statorassemblies and causes the rotor assemblies to rotate a main engine shaftfor energy extraction. The air is then exhausted through a propulsionnozzle disposed in the exhaust section 160 to provide additional forwardthrust.

FIG. 2 is a partial cross-sectional side view of a turbine section of anengine, such as the turbine section 150 of engine 100 of FIG. 1 inaccordance with an exemplary embodiment. The turbine section 150includes a turbine stator 200 and a turbine rotor 250 surrounded by ashroud 210 defining a gas flow path through which hot, combusted airfrom an upstream compressor section (e.g. compressor section 130 of FIG.1 ) is directed. The cylindrical shroud 210 is disposed concentric tothe rotor 250 to optimize aerodynamic efficiency and forms a radial gap(i.e., blade running clearance) 270 with an outermost diameter of therotor 250. Although only one turbine stator 200 and one turbine rotor250 are shown, such stators 200 and rotors 250 are typically arranged inalternating axially spaced, circumferential rows. As used herein, theterm “axial” refers to a direction generally parallel to the enginecenterline, while the term “radial” refers to a direction generallyperpendicular to the engine centerline.

The rotor 250 generally includes rotor blades 260 (one of which isshown) mounted on a rotor disc (not shown), which in turn is coupled toan engine shaft (not shown). The turbine stator 200 directs the airtoward the turbine rotor 250. The air impinges upon rotor blades 260 ofthe turbine rotor 250, thereby driving the turbine rotor 250 for powerextraction. To allow the turbine section 150 to operate at desirableelevated temperatures, certain components are coating with the EBC/TBCcoatings of the present disclosure, such as the shroud or nozzles.

Ceramic Substrate Materials

As noted above, various hot section components as illustrated in FIG. 2may be formed of a silicon nitride or silicon carbide fiber/siliconcarbide matrix composite material. In one example, as generally known inthe art, a SiC-SiC ceramic matrix composite material may include a SiCfiber-bonded ceramic or a SiC fiber-bonded ceramic having a gradedstructure, for example. Regarding the SiC fiber-bonded ceramic, such amaterial may generally include inorganic fibers having mainly a sinteredSiC structure, each of which contains 0.01-1 wt. % of oxygen (O) and atleast one or more metal atoms of metal atoms in Groups 2A, 3A, and 3B,and a 1-100 nm interfacial layer containing carbon (C) as a maincomponent formed between the fibers. Further, the SiC fiber-bondedceramic having a graded structure may generally include a matrix, thematrix including inorganic fibers having mainly a sintered SiC structurecontaining 0.01-1 wt. % of oxygen (O) and at least one or more metalatoms of metal atoms in Groups 2A, 3A, and 3B, and a 1-100 nminterfacial layer containing carbon (C) and/or boron nitride (BN) as amain component formed between the fibers, a surface portion having aceramic structure including mainly SiC and being formed on at least partof the surface of the matrix, a boundary portion interposed between thesurface portion and the matrix and having a graded structure thatchanges from the structure of the matrix to the structure of the surfaceportion gradually and continuously.

These SiC—SiC materials include a volume fraction of about 90% or moreof SiC-based fibers. Such materials have high fracture toughness and areinsensitive to defects. The fiber material constituting the SiCfiber-bonded ceramic is mainly inorganic fibers that include a sinteringstructure containing mainly SiC and/or SiCN, contain about 0.01-1 wt. %of oxygen (O) and at least one metal atom selected from the groupincluding metal atoms in Groups 2A, 3A, and 3B, and are bonded veryclose to the closest-packed structure. The inorganic fibers including asintered SiC structure include mainly a sintered polycrystalline n-SiCstructure, or include crystalline particulates of β-SiC and C. In aregion containing a fine crystal of carbon (C) and/or an extremely smallamount of oxygen (O), where β-SiC crystal grains sinter together withoutgrain boundary second phase interposed therebetween, a strong bondbetween SiC crystals can be obtained.

FIG. 3 illustrates, in cross section, the surface of an exemplaryturbine component substrate 300. FIG. 3 is a cross-sectional view ofsubstrate 300 formed of a SiC-SiC material as described above upon whichis to be disposed a protective coating system in accordance with anexemplary embodiment of the present disclosure. As shown in FIG. 3 , thesubstrate 300 has a generally irregular or “wavy” outer surface 310,including “pits” and “valleys,” upon which the protective coating systemis to be disposed, and which may be formed by the woven fibers of theceramic matric composite. The irregular surface includes deviations(+/−) from planar of several mils, such as about 1 mil to about 5 mils.The outer surface 310 may also have larger defects beyond theillustrated irregular surface. These larger defects may includedeviations from planarity of about 30 mils to about 50 mils or greater,in some instances.

Rare Earth Disilicate EBC Layer

Referring now to FIG. 4 , disposed over the substrate surface 310 is EBCcoating layer 320. EBC coating layer 320 is formed, for example, to athickness from 10 mils to about 100 mils, for example from about 20 milsto about 50 mils. The EBC coating layer 320 is provided as anenvironmental barrier coating to enhance the operating capabilities of agas turbine engine component fabricated with the SiC-SiC substrate asdescribed above. For example, in one embodiment, the EBC coating layeris provided for protection in high temperature and high pressureenvironments of the turbine engine, which can cause oxidation andeventual erosion of silicon based ceramics. In particular embodiments,the addition of EBC coating layer 320, in part, functions to raise thesuitable operating temperature of a component formed with substrate 300to about 2600° F. to about 2800° F. or greater.

The EBC coating layer 320 may generally include a rare-earth disilicatematerial. As known in the art, rare earth elements include, amongvarious others, strontium, lanthanum, yttrium, scandium, and others. Forsome embodiments, disilicate of yttrium and scandium are particularlysuitable. The EBC coating layer 320 may be applied over the substrate300 coating layer using any known methods. These methods include, butare not limited to, plasma spraying, physical vapor deposition (PVD),and electron beam physical vapor deposition (EB-PVD), and dipping.

Conversion of Disilicate to Monosilicate

As initially noted, the approach employed in the present disclosureutilizes the rare earth disilicate EBC coating, just described, that isdisposed onto the substrate, and then subjected to thermal or carbonmonoxide-based processing to decompose an outer surface of the rareearth disilicate coating to form a rare earth monosilicate porous outerlayer. This porous outer layer may then be infiltrated with a metal saltsolution or a metal oxide nanoparticle suspension. The disilicatedecomposes to the more stable monosilicate phase above about 2400° F.surface temperature in a combustion environment. This monosilicatetransformation will occur normally in service, but it can be produced byother methods in more controlled manners. These other methods result ina porous surface structure that can be further treated to form morestable high temperature oxides or silicides through the impregnation ofdifferent metal nitrates, carbonates or oxides that can be reacted withthe surface in a step separate from the coating sintering where there isrisk of different phases to cause problems with the densification of thecoating. The advantage of this being that high temperature phases can beachieved at temperatures low enough to prevent continued sintering ofthe EBC. Moreover, because this is a porous layer that is formed in astable condition from the starting EBC, there is a chemical transitionthat occurs between the dense disilicate EBC and the porous monosilicatesurface.

FIG. 5 illustrates the substrate 300 having the disilicate EBC coating320 thereon having been partially converted to porous monosilicate layer330. The thickness of the porous layer 330 may be from about 5 micronsto about 25 microns, such as about 10 microns to about 20 microns, andwill have consumed some of the disilicate layer, resulting in this layerbeing commensurately thinner than previously described. The pores may beof various shapes and sizes. In some embodiments, the porosity ofcoating layer 330 may be from about 10% to about 70% (the percentageindicates the amount, by volume, of void space as a result of thepresence of pores). In other embodiments, the porosity may be from about25% to about 50%. The number and distribution of pores may besubstantially equivalent throughout the thickness of layer 330. In otherembodiments, processing may be provided such that there is a porositygradient within layer 330. For example, a greater or lesser degree ofporosity may be provided in areas of layer 330 that are relative closerto layer 320, whereas a lesser or greater degree of porosity may beprovided in areas of layer 330 that are relatively further from layer320. As known in the art, greater porosity provides greater thermalbarrier capabilities, but renders the material less stable. Higherporosity will also likely compromise the mechanical properties of thecoating. As such, in one embodiment, a relatively lesser porosity (forexample from about 10% to about 40%) is provided in areas of layer 330that are relatively closer to layer 320, and a relatively greaterporosity (for example from about 40% to about 70%) is provided in areasof layer 330 that are relatively further from layer 320.

The conversion of the disilicate to the monosilicate may be accomplishedin several manners. Using either manner, this surface treatment willimprove chemical stability of the silicon nitride or SiC/SiC componentEBC coatings at high temperatures, provide a mechanically compliantlayer that reduced CTE mismatch thermal stresses, reduce the coatingthermal conductivity coefficient for all turbine components, reducingsecondary cooling loads, and provide abradability to improve blade tiprub tolerance for shrouds. The treatment will improve engine performanceby increasing engine efficiency. Being able to increase cycletemperatures allows for higher thermodynamic efficiency. The reducedcooling load lower cooling requirements reducing parasitic cooling loadsand potentially reduce engine core size.

For example, in one embodiment, a thermal treatment is employed. Asinitially noted, the disilicate decomposes to the more stablemonosilicate phase above about 2400° F. surface temperature in acombustion environment. Accordingly, for the thermal treatment, thecoated turbine component may be exposed to temperatures in excess of2400° F. in a reducing/oxygen-poor environment for a period of time.Moreover, depending on the degree of porosity desired to be formed fromthe layer 320, two or more cycles of heating may be employed, ranging invarious times and temperatures. Thermal treatment can be by the use of aflame, torch, or in more controlled environment such as a furnace.Higher CO concentrations and higher temperatures increase thetransformation rates. Transformation rates dictate morphology and porestructures (for example highly-porous, finger-like structures onsurface, depth of transformation, etc.) FIG. 6 is a cross-sectionalmicrograph of an EBC layer 320 having been thermally treated asdescribed above to form porous monosilicate layer 330. As can be seen,the porosity increases from the interface with layer 320 to the outersurface.

In another embodiment, conversion of the disilicate to the monosilicatemay be accomplished using a chemical-thermal treatment with exposure tocarbon monoxide. Referring to FIG. 8 , the disilicate EBC coated part(block 350) is exposed to temperatures 1400° C., preferably above 1500°C., or higher in a flow of carbon monoxide 325 (block 351). Above thistemperature, a carbothermal reduction process will convert thedisilicate coating to a monosilicate by the off gassing of SiO(g) andcarbon dioxide 327. Laboratory observation has shown that the resultingsurface of the coating is porous monosilicate layer that is intimatelybonded to the underlying, unaffected dislicate EBC with a compositionalgradient. The depth of the transformation can be controlled by time,temperature, CO concentration, and gas flow rate. For example, it hasbeen observed that when the CO concentration is estimated to be in therage of 1000 ppm (or 0.1%) at temperatures above 1550° C. thetransformation occurred in about 30 minutes. Because this is a layerthat is formed from the starting EBC, there is chemical and transitionthat occurs between the dense disilicate EBC and the porous monosilicateTBC surface. FIG. 7 is a cross-sectional micrograph of an EBC layer 320having been chemically-thermally treated as described above to formporous monosilicate layer 330. As can be seen, the porosity increasesfrom the interface with layer 320 to the outer surface.

Infiltration of Metal Salts, Nitrates, Carbonates and/or Oxides

As initially noted, the previously-described forming of the porousmonosilicate layer 330 results in a porous surface structure that can befurther treated to form more stable high temperature oxides or silicidesthrough the impregnation of different metal nitrates, carbonates, and/oroxides that can be reacted with the surface in a step separate from thecoating sintering where there is risk of different phases to causeproblems with the densification of the coating. The advantage of thisbeing that high temperature phases can be achieved at temperatures lowenough to prevent continued sintering of the EBC. Depending or whatmetal ions are introduced and to what degree, temperature stability canbe further increased, or the coating can potentially be given anincreased CMAS resistance. The use of this method on a coating alreadytransformed to monosilicate with a porous structure reduces the risk ofspallation due to any thermo-mechanical stresses associated with thenewly formed phases caused by the metal ion additions or any stressesassociate with thermal gradient during engine operation or enginetransients. The metals of this infiltration layer may include, withoutlimitation, materials including aluminum, zirconium, titanium, yttrium,hafnium, tantalum, ytterbium, strontium, and the like, depending on thedesired function. The infiltration may be accomplished using a metalsalt solution or a metal oxide nanoparticle suspension, as describedbelow.

In accordance with one embodiment, the porous surface may be infiltratedwith a metal salt solution with the desired metal ion (e.g. nitrate,carbonate, sulphate, chloride). The impregnated coating is dried in airand the component is calcined in air at a temperature above thedecomposition temperature of the metal salt, resulting the metal oxide.The component can then be treated at high temperature to sinter or reactthe newly formed metal oxide to the base silicate coating. The amount ofdesired metal oxide added is controlled by the concentration of the saltsolution.

In accordance with another embodiment, the porous surface may beinfiltrated with a nano metal oxide suspension. For this embodiment, asuspension with nano metal oxide particles, surfactants, and binders (toensure wetting) is prepared. The suspension is then applied to thetreated coating surface. The impregnated coating dries, and then theorganic binders and surfactants are burned out in air at anappropriately high temperature, depending on the exact binders andsurfactants utilized. Finally, the coted component is heat treated athigh temperature to sinter or react the nano particles to the coating.The amount of desired metal oxide added is controlled by theconcentration of the solid loading of the nano particle suspension.

In either embodiment, all post-infiltration heat treatment temperatureswill depend on the desired metal ion that will be introduced and theirdecomposition temperature, the reactivity/diffusivity of the resultingoxide with the monosilicate coating. For the nano metal oxide approach,the difference in the sintering temperature of the nanoparticlescompared to the sintering temperature of the coating may be exploited toprefer the formation of a metal oxide surface over the monosilicate orthe reaction with the monosilicate forming silicate with more than onemetal ion. Temperatures between about 1000° C. and about 1500° C. areexemplary.

Further, in either embodiment, application of the salt solution or nanometal oxide suspension may be accomplished using, for example,spin-coating, dip-coating, spraying, roll coating, and others. Inembodiments wherein there is a gradient in pore size, with smaller poresbeing located near layer 320, the metal oxide material may onlyinfiltrate a certain portion of the depth of layer 330 (namely into thepores thereof). For example, in some embodiments, the metal oxide mayinfiltrate 70% or less of the thickness of layer 330, 50% of thethickness or less, or even 30% of the thickness or less. The amount ofmaterial of the metal oxide deposited thus depends on the thickness anddesired fill properties of layer 330. Exact thickness for a givenembodiment will ultimately be determined by the skilled artisan, but maygenerally be from about 5 mils to about 20 mils, as initially deposited.Infiltration of the metal oxide material into layer 330 may beaccomplished in a variety of manners, including for example capillaryaction or an applied vacuum.

Infiltration Example #1—Salt Solution

An aqueous solution of hafnium chloride (HfCl₄) is prepared. Themonosilicate layer surface is wet with this solution, dried, andcalcined at 250° C. to decompose the salt to hafnium oxide. To increasethe hafnium oxide loading, this step is repeated. A final heat treatmentat >1250° C. is done to react the HfO₂ with the monosilicate resultingin a rare earth silicate (some of the oxide reacts to form the silicateand some remains as the oxide). The resulting coating layer isillustrated in FIG. 9 , in micrograph.

Infiltration Example #2—Nano Metal Oxide Suspension

A suspension with a high concentration of nano HfO₂ particles isprepared with and acrylic binder and surfactant to disperse the HfO₂particles and ensure wetting of the monosilicate surface. The suspensionis used to impregnate the monosilciate coating and is dried at 150° C.Any organic are then burned out in air at 300° C. A final heat treatmentis done at >1250° C. to react the HfO₂ with the monosilicate resultingin a rare earth silicate. The resulting coating layer is illustrated inFIG. 10 , in micrograph.

Accordingly, protective coating systems for gas turbine engineapplications and methods for fabricating such protective coating systemshave been provided. The disclosed embodiments beneficially provide anovel approach to creating an EBC/TBC coating for SiN and SiC-SiCsubstrate materials to allow the use temperature to be increased fromabout 2600° F. to about 2800° F.

While at least one exemplary embodiment has been presented in theforegoing detailed description of the invention, it should beappreciated that a vast number of variations exist. It should also beappreciated that the exemplary embodiment or exemplary embodiments areonly examples, and are not intended to limit the scope, applicability,or configuration of the invention in any way. Rather, the foregoingdetailed description will provide those skilled in the art with aconvenient road map for implementing an exemplary embodiment of theinvention, it being understood that various changes may be made in thefunction and arrangement of elements described in an exemplaryembodiment without departing from the scope of the invention as setforth in the appended claims and their legal equivalents.

what is claimed is:
 1. A protective coating system comprising: a turbineengine component substrate formed of a ceramic matrix compositematerial; an environmental barrier coating layer comprising a rare earthdisilicate material formed directly on the substrate; and a thermalbarrier coating layer comprising a porous rare earth monosilicatematerial having a metal silicate material infiltrated within at least aportion of the pores formed directly on the environmental barriercoating layer.
 2. The protective coating system of claim 1, wherein theceramic matrix composite material is a silicon carbide-silicon carbide(SiC—SiC) material or a silicon nitride (Si₃N₄) material.
 3. Theprotective coating system of claim 1, wherein the metal silicatematerial comprises a metal element selected from the group consistingof: aluminum, zirconium, titanium, yttrium, hafnium, tantalum,ytterbium, and strontium.
 4. The protective coating system of claim 1,wherein the environmental barrier coating layer has a thickness of about10 mils to about 100 mils.
 5. The protective coating system of claim 1,wherein the thermal barrier coating layer comprises a plurality of poressuch that it comprises a porosity of about 10% to about 70% by volume.6. The protective coating system of claim 5, wherein the porosity has agradient within the second coating layer.
 7. The protective coatingsystem of claim 1, wherein the metal silicate material infiltrates thethermal barrier coating layer to a depth of about 70% or less of a totalthickness of the thermal barrier coating layer, thereby creating aporosity gradient in the thermal barrier coating layer.
 8. A method ofapplying a protective coating to a substrate comprises the steps of:providing a turbine engine component substrate formed of a ceramicmatrix composite material; forming an environmental barrier coatinglayer comprising a rare earth disilicate material directly on thesubstrate; treating an outer surface of the environmental barriercoating layer to form a thermal barrier coating layer comprising aporous rare earth monociliate material directly on the environmentalbarrier coating layer; and infiltrating at least a portion of the poreswith a metal solution or suspension.
 9. The method of claim 8, whereinthe ceramic matrix composite material is a silicon carbide-siliconcarbide (SiC—SiC) material or a silicon nitride (Si₃N₄) material. 10.The method of claim 8, wherein the metal solution or suspensioncomprises a metal element selected from the group consisting of:aluminum, zirconium, titanium, yttrium, hafnium, tantalum, ytterbium,and strontium.
 11. The method of claim 10, wherein the metal solution orsuspension comprises a metal nano oxide or a metal salt, nitrate,carbonate or oxide.
 12. The method of claim 8, wherein the environmentalbarrier coating layer is formed to have a thickness of about 10 mils toabout 100 mils.
 13. The method of claim 8, wherein the step of treatingthe outer surface is performed using a thermal process or achemical-thermal process.
 14. The method of claim 8, further comprisingheat-treating the substrate after infiltrating the pores with the metalsolution to form a metal silicate material within the pores.
 15. Themethod of claim 8, wherein infiltrating the metal solution or suspensionis performed to infiltrate the thermal barrier coating layer to a depthof about 70% or less of a total thickness of the thermal barrier coatinglayer, thereby creating a porosity gradient in the thermal barriercoating layer.